Location-specific slurry based coatings for internally-cooled component and process therefor

ABSTRACT

A method of coating a component including aluminizing an array of internal passageways within the component; and chromizing a portion of the array of internal passageways within the component. A component, including an airfoil having an array of aluminized internal passageways, the array of aluminized internal passageways chromized up to a demarcation.

BACKGROUND

The present disclosure relates to coating and, more particularly, toslurry coating compositions in which the properties of the coating aretailored to resist local conditions within an internally-cooledcomponent.

Gas turbine engines typically include a compressor section to pressurizeairflow, a combustor section to burn a hydrocarbon fuel in the presenceof the pressurized air, and a turbine section to extract energy from theresultant combustion gases. Gas path components, such as turbine blades,often include airfoil cooling that may be accomplished by external filmcooling, internal air impingement and forced convection eitherseparately, or in combination.

The internal cavities include internal passages to direct the passage ofthe cooling air. As gas turbine temperatures have increased, thegeometries of these cooling passages have become progressively morecircuitous and complex. Such internal passages are often coated due todeposit-induced hot corrosion as well as high temperature oxidation.Coatings currently in production largely address one of thesedegradation modes but may be less than effective when used inenvironments for which they are not optimized.

SUMMARY

A method of coating a component according to one disclosed non-limitingembodiment of the present disclosure includes aluminizing an array ofinternal passageways within the component, and chromizing a portion ofthe array of internal passageways within the component after aluminizingthe array of internal passageways.

A further aspect of the present disclosure includes chromizing up to ademarcation.

A further aspect of the present disclosure includes that the demarcationis located at 20%-30% of an airfoil span of the component.

A further aspect of the present disclosure includes that the demarcationis at 25% of an airfoil span of the component.

A further aspect of the present disclosure includes that the demarcationis at an airfoil span location where local temperatures exceedapproximately 1600 degree F.

A further aspect of the present disclosure includes applying a chromiumslurry into the array of internal passageways within the component.

A further aspect of the present disclosure includes that the chromiumslurry has a viscosity of 100-200 cp.

A further aspect of the present disclosure includes applying a thermalbarrier coat atop a bond coat on an external surface of the component.

A component, according to one disclosed non-limiting embodiment of thepresent disclosure, includes a substrate having an array of aluminizedinternal passageways within the component, and a chromium-enrichedcoating within a portion of the array of internal passageways.

A further aspect of the present disclosure includes a superalloy.

A further aspect of the present disclosure includes that thechromium-enriched coating within the portion of the array of internalpassageways is located at an inboard section of the component.

A further aspect of the present disclosure includes that thechromium-enriched coating within the portion of the array of internalpassageways is located up to a demarcation that is located at 20%-30% ofan airfoil span of the component.

A further aspect of the present disclosure includes that thechromium-enriched coating within the portion of the array of internalpassageways is located up to a demarcation that is located at 25% of anairfoil span of the component.

A further aspect of the present disclosure includes that thechromium-enriched coating within the portion of the array of internalpassageways is located up to an airfoil span location where localtemperatures exceed approximately 1600 degree F.

A further aspect of the present disclosure includes that thechromium-enriched coating is 10-50 microns thick.

A component according to one disclosed non-limiting embodiment of thepresent disclosure includes an airfoil having an array of aluminizedinternal passageways, the array of aluminized internal passagewayschromized up to a demarcation.

A further aspect of the present disclosure includes that the demarcationis located at 20%-30% of a span of the airfoil.

A further aspect of the present disclosure includes that the demarcationis located at 25% of a span of the airfoil.

A further aspect of the present disclosure includes that the demarcationis located is at an airfoil span location where local temperaturesexceed approximately 1600 degree F.

A further aspect of the present disclosure includes that the componentis a turbine blade.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, that the followingdescription and drawings are intended to be exemplary in nature andnon-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiments. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of an example gas turbine enginearchitecture.

FIG. 2 is an enlarged schematic cross-section of an engine turbinesection.

FIG. 3 is a perspective view of an airfoil as an example component foruse with a coating method showing an example internal architecture.

FIG. 4 is a block diagram representing a method of coating an array ofinternal passageways of a component.

FIG. 5 is a micrograph of an aluminide coating before any chromizingwithin a portion of an array of internal passageways prior tochromizing.

FIG. 6 is a micrograph of a portion of the array of internal passagewaysin which the aluminide coating that has been chromized.

FIG. 7 is a micrograph of a component processed with the method ofcoating the array of internal passageways after 1000 hours in a hotcorrosion test at 1350 F.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplarygas turbine engine 20 is disclosed herein as a two-spool turbo fan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flowpath and along a core flowpath for compression bythe compressor section 24, communication into the combustor section 26,then expansion through the turbine section 28. Although depicted as aturbofan in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith turbofans as the teachings may be applied to other types of turbineengine architectures such as low bypass turbofans, turbojets,turboshafts, three-spool (plus fan) turbofans and other non-gas turbinecomponents.

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation an engine central longitudinal axis “A”. The lowspool 30 generally includes an inner shaft 40 that interconnects a fan42, a low pressure compressor (“LPC”) 44 and a low pressure turbine(“LPT”) 46. The inner shaft 40 drives the fan 42 directly, or through ageared architecture 48 at a lower speed than the low spool 30. Anexemplary reduction transmission is an epicyclic transmission, namely aplanetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor (“HPC”) 52 and high pressure turbine (“HPT”) 54. Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate the engine central longitudinal axis “A”, which iscollinear with their longitudinal axes.

Core airflow is compressed by the LPC 44, then the HPC 52, mixed withthe fuel and burned in the combustor 56, then expanded over the HPT 54,then the LPT 46. The turbines 54, 46 rotationally drive the respectivehigh spool 32 and low spool 30 in response to the expansion. The mainengine shafts 40, 50 are supported at a plurality of points by bearingstructures 38 within the static structure 36.

With reference to FIG. 2, an enlarged schematic view of a portion of theturbine section 28 is shown by way of example; however, other enginesections will also benefit herefrom. A shroud assembly 60 within theengine case structure 36 supports a blade outer air seal (BOAS) assembly62 with a multiple of circumferentially distributed BOAS 64 proximate toa rotor assembly 66 (one schematically shown).

The shroud assembly 60 and the BOAS assembly 62 are axially disposedbetween a forward stationary vane ring 68 and an aft stationary vanering 70. Each vane ring 68, 70 includes an array of vanes 72, 74 thatextend between a respective inner vane platform 76, 78 and an outer vaneplatform 80, 82. The outer vane platforms 80, 82 are attached to theengine case structure 36.

The rotor assembly 66 includes an array of blades 84 circumferentiallydisposed around a disk 86. Each blade 84 includes a root 88, a platform90 and an airfoil 92 (also shown in FIG. 3). The blade roots 88 arereceived within a rim 94 of the disk 86 and the airfoils 92 extendradially outward such that a tip 96 of each airfoil 92 is closest to theblade outer air seal (BOAS) assembly 62. The platform 90 separates a gaspath side inclusive of the airfoil 92 and a non-gas path side inclusiveof the root 88.

With reference to FIG. 3, the platform 90 generally separates the root88 and the airfoil 92 to define an inner boundary of a gas path. Theairfoil 92 defines a blade chord between a leading edge 98, which mayinclude various forward and/or aft sweep configurations, and a trailingedge 100. A first sidewall 102 that may be convex to define a suctionside, and a second sidewall 104 that may be concave to define a pressureside are joined at the leading edge 98 and at the axially spacedtrailing edge 100. The tip 96 extends between the sidewalls 102, 104opposite the platform 90. It should be appreciated that the tip 96 mayinclude a recessed portion.

To resist the high temperature stress environment in the gas path of aturbine engine, each blade 84 may be formed by casting. It should beappreciated that although a blade 84 with an array of internalpassageways 110 (shown schematically) will be described and illustratedin detail, other hot section components including, but not limited to,vanes, turbine shrouds, end walls and other components will also benefitfrom the teachings herein.

The external airfoil surface may be protected by a protective coatingthat overlies and contacts the external airfoil surface. Such coatingsmay be of the MCrAIX type. The terminology “MCrAIX” is a shorthand termof art for a variety of families of overlay protective layers that maybe employed as environmental coatings or bond coats in thermal barriercoating systems. In this, and other forms, M refers to nickel, cobalt,iron, and combinations thereof. In some of these protective coatings,the chromium may be omitted. The X denotes elements such as hafnium,zirconium, yttrium, tantalum, rhenium, ruthenium, palladium, platinum,silicon, titanium, boron, carbon, and combinations thereof. Specificcompositions are known in the art. Optionally, a ceramic layer overliesand contacts the protective layer. The ceramic layer is preferablyyttria-stabilized zirconia, which is a zirconium oxide. Other operableceramic materials may be used as well. Often, when there is no ceramiclayer present, the protective layer is termed an “environmentalcoating.” When there is a ceramic layer present, the protective layer isoften referred to as a “bond coat”.

The array of internal passageways 110 generally includes one or morefeed passages 112 that communicate airflow into a trailing edge cavity114 within the blade 84. It should be appreciated that the array ofinternal passageways 110 may be of various geometries, numbers andconfigurations and the feed passage 112 in this embodiment is the aftmost passage that communicates cooling air to the trailing edge cavity114. The feed passage 112 generally receives cooling flow through atleast one inlet 116 within a base 118 of the root 88.

The trailing edge cavity 114 may include a multiple of trailing edgecavity features 120 that result in a circuitous and complex coolingairflow path. It should be appreciated that although particular featuresare delineated within certain general areas, the features may beotherwise arranged or intermingled and still not depart from thedisclosure herein.

The array of internal passageways 110 are generally present in variousgas turbine components, such as the example blade 84, to allow for thepassage of cooling air. As gas turbine temperatures have increased, thegeometries of these cooling passages have become progressively morecircuitous and complex. These internal passages 110, as well as otherportions of the workpiece, are often coated with a metallic coatingapplied via a diffusion chromizing process to prevent hot corrosion.Generally, components are placed in a retort for distillation,Cr-containing vapor species are generated and supplied to the surface ofthe components via gas phase transport, and a Cr-rich coating is formed.Although effective, the vapor phase chromizing process may suffer froman inability to achieve sufficient coverage and Cr content on somecomponents, particular High Pressure Turbine (HPT) blades that areeither very large or that contain very complex internal passages.

The example component workpiece, such as the blade 84, is typicallymanufactured of a nickel-base alloy, and more preferably of anickel-base superalloy. A nickel-base alloy has more nickel than anyother element, and a nickel-base superalloy is a nickel-base alloy thatis strengthened by the precipitation of gamma prime or a related phase.The component, and thence a substrate and the internal passagewaysthereof, are thus of nickel-base alloy, and more preferably are anickel-base superalloy.

In this embodiment, the array of internal passageways 110 within theblade 84 or other workpiece includes an aluminized section 130 and achromized section 132. That is, the entirety of thepreviously-aluminized array of internal passageways 110 are chromizedonly up to a demarcation 140 such that only the chromized section 132has a relatively thick (e.g., 1.5-3 mil) Cr-rich coating 144. In thisembodiment, the demarcation 140 may be located at 20%-30% the span ofthe airfoil 92 and more specifically at a 25% span location of theairfoil 92. That is, the previously-aluminized array of internalpassageways 110 is chromized up to the demarcation 140. Aluminides areprotective in high temperature oxidation and less so in hot corrosion,while the reverse is true for chromizing.

Alternatively, the demarcation 140 may be defined by a temperature rangeexperienced by the airfoil 92 during operation. For example, thedemarcation 140 may be defined at the location on the airfoil 92 wherelocal temperatures exceed approximately 1600 degree F. Typically, hightemperature oxidation occurs around 1700 F-1800 F which appears in theoutboard section of the airfoil 92 such that the aluminides in thealuminized section 130 forms an aluminum oxide scale this is resistantto oxidation damage at high temperatures while the relatively coolerchromized section 132 is resistant to contaminates at relatively coolerinboard sections. That is, the aluminized section 130 and the chromizedsection 132 provide effective resistance to oxidation anddeposit-induced hot corrosion, respectively when flown in pollutedenvironments.

With reference to FIG. 4, one disclosed non-limiting embodiment of amethod 200 for applying a location-specific coating is illustrated. Thesteps or actions described with respect to the method 200, can beemployed with additional steps or other processes as desired. Although aturbine blade 84 is illustrated as the article in the disclosedembodiment, the article may be a metallic article formed of a base alloysubstrate. For example, the base alloy substrate is a superalloy. In afurther example, the superalloy is a nickel-base alloy. In a stillfurther example, the base alloy is a low-chromium superalloy, such as asuperalloy having less than approximately 12% by weight chromium. Itshould be appreciated that although the blade 84 is illustrated in thedisclose embodiment, any such component desired to have internal hotcorrosion protection and high temperature oxidation resistance willbenefit herefrom.

Initially, the first step of the process 200 is to aluminize (step 202;FIG. 5) the component internals, e.g., the array of internal passageways110. The aluminizing may be accomplished by chemical vapor deposition(CVD), above the pack, or slurry processes. The aluminizing alone willbe the mitigation strategy in the outer portions outboard of thedemarcation 140 where high temperature oxidation damage typicallydominates. In addition, the externals may be aluminized as well.

The component may be exposed to an aluminum source material such as achromium-aluminum alloy in the presence of an activator, such as ahalide activator, and a cover gas such as argon or hydrogen, at atemperature of between approximately 1900 F-2100 F (1037 C-1148 C) for atime of between 1-6 hours. The halide activator can be, for instance, afluoride or a chloride in a solid or gaseous form.

In another example aluminizing process, the aluminum or aluminum sourceand activator can be combined with another material such as aluminumoxide powder prior to being deposited on the component. The othermaterial can enhance the deposition of aluminum or the properties of thedeposited aluminum. Then, the component is exposed to the aluminum oraluminum mixture at a temperature of less than approximately 1500 F (815C) for a time of between approximately 2-4 hours.

In another example aluminizing process, the aluminum or aluminum sourceand activator can be combined with another material such as aluminumoxide powder prior to being deposited on component to enhance thedeposition of aluminum or the properties of the deposited aluminum.Then, the component is exposed to the aluminum or aluminum mixture at atemperature of between 1500 F-1900 F (815 C-1037 C) for a time ofbetween 2-8 hours. In another example process, the aluminum or aluminumsource and activator can be combined with another material such asaluminum oxide powder and a binder to form a slurry prior to beingdeposited on component. The other material can enhance the deposition ofaluminum or the properties of the deposited aluminum. Then, thecomponent is exposed to the aluminum or aluminum mixture at atemperature of between 1500 F-1900 F (815 C-1037 C) for a time between2-8 hours.

Any of the aluminizing processes described above can be repeated toincorporate additional aluminum multiple times. Additionally, any of theprocesses described above can form an aluminum coating that is betweenapproximately 0.5 and 3.0 thousandth of an inch (0.01 to 0.08 mm) thick,for example.

After aluminizing to provide an aluminized coating 142 in the array ofinternal passageways 110, the component may be masked (step 204). Themask may be performed via plugging/blocking of the openings to the arrayof internal passageways 110. Other particular external areas such as theroot, underplatform, and airfoil external surfaces may also be masked bysacrificial coating, taping, mechanical fixturing/masking or the like.

Next, the component is chromized (step 206; FIG. 6) up to thedemarcation 140. For example, only an inboard portion, e.g., from theroot 88, platform 90, and the airfoil 92, of the component is dippedinto a tank containing a chromizing slurry so that the slurry is appliedto the internal surfaces up to the desired demarcation 140 of theairfoil 92. The chromium slurry, for example, can be flowed through thecomponent to achieve coverage on complex geometries, here, the array ofinternal passageways 110. The chromium slurry may alternatively beapplied to the component, for example, by pouring, injecting, orotherwise flowing the slurry into the array of internal passageways 110.In another disclosed non-limiting embodiment, a portion of the componentis dipped therein. Alternately, the chromium slurry is applied via othercarriers, devices, and/or methods. The resultant chromium enricheddiffusion coating 144 (FIG. 6; the aluminide after being chromized) interms of microstructure and thickness is similar to the aluminidecoating 142 (FIG. 5; before any chromizing) but a substantial amount ofthe aluminum in the coating is replaced with chromium. The vaporaluminide shown in FIG. 5 may be deposited by a high temperature, lowactivity aluminizing process. An outer layer (dimension B) which may beabout 0.9 mil, is the beta-NiAl phase, which grows outward from theoriginal alloy surface; the remainder of the coating is theinter-diffusion zone (IDZ). The entire chromium enriched diffusioncoating 144 (FIG. 6; dimension B+X) in one example is about 1.6 mil. Theouter layer of the coating (dimension B; FIG. 6) was formerly beta-NiAlthat was converted to either Cr-rich gamma-Ni solid solution orgamma+alpha-Cr upon chromizing.

The chromium slurry may include a mixture of chromium powder, chromiumChloride (CrCl3) particles as an activator, and, optionally, an organicbinder. While using a slurry with no oxide filler is an option, otherembodiments may utilize either Al2O3 or SiO2 in this process. SiO2 hasthe benefit of being removable via autoclave with a basic solution ofKOH or NaOH, which may be desirable for the array of internalpassageways 110. Other slurry coatings contain aluminum oxide filler,but the presence of such a filler in a coating slurry that is used tocoat internal surfaces such as the array of internal passageways 110 maybe a primary cause of undesirable obstruction and/or flow disturbanceswithin the array of internal passageways 110. The chromium (Cr) slurry,in one example in terms of weight percentages, includes 48.5-68% byweight chromium powder, 0.9-3.4% by weight chromium Chloride (CrCl3)particles, and 30-50% by weight organic binder. The resultant chromium(Cr) slurry forms a low-viscosity fluid capable of being flowed throughinternal passages. In one example, the slurry has a viscosity of 100-200cp. Any operable organic binder may be used. Examples include, but arenot limited to, B4 (n-propyl bromide-based organic binder such as thatfrom Akron Paint and Varnish) and Klucel H (hydroxypropyl cellulose),and mixtures thereof. Other organic binders such as a water basedorganic binder may alternatively be utilized.

The chromium (Cr) slurry, in another example without an organic binderin terms of weight percentages, includes 97% by weight chromium powderand 0.03% by weight chromium Chloride (CrCl3) particles.

Next, the excess chromium slurry is drained away (step 208). Simplyallowing the relatively viscous chromium slurry to flow out of theinternal passageways 110 may perform such draining.

The chromium slurry is then dried to drive off the organic binder (step210). The drying evaporates the flowable carrier component of theorganic binder (e.g., flowable organic solvents and water) of thechromium slurry, leaving the organic binder that binds the particlestogether. Driving off the organic binder is performed at a relativelylow temperature for short periods of time. In one example, drying of thebinder is performed at 200 F (93 C) for 1 hour. Alternatively, thedrying could be performed at room temperature given a commensurategreater time period. The applying, draining and drying steps may also berepeated multiple times to achieve a desired thickness and/or coverage.

Next, the component is heat treated (step 212). In one example, heattreat may be accomplished at a temperature of from 1600 F (871 C) to2100 F (1149 C) most preferably from 1925 F (1052 C) to 2000 F (1093 C),for a time of 4-8 hours, preferably, 5-6 hours. The heat treating may beperformed in an inert (e.g., argon) or reducing (e.g., hydrogen)atmosphere. In the case of the inert atmosphere, the atmosphere islargely free of oxygen and oxygen-containing species such as watervapor.

The heat treat allows, through a mechanism involving the reaction of theCr powder with the activator, gas phase transport of Cr-containingspecies to the component surface, and subsequent diffusion of Cr intothe parent material, the formation of a coating that, in one disclosednon-limiting embodiment, is 10-50 microns thick chromium-enriched singlephase y face centered cubic Ni-based solid solution layer that preventshot corrosion. There may be some transition zone to which the Cr-richvapors travel, but the outboard portion of the blade will remainaluminized with minimum Cr deposition. If the aluminizing isaccomplished via a slurry process, the possibility exists to onlyconduct one heat treatment.

After the heat treatment the “spent” slurry is removed (step 214). Thereis essentially a friable crust of Cr-containing powder on the array ofinternal passageways 110 after the heat treatment, and this is to beremoved. In one example, warm hydrogen chloride (HCl) may be utilized todissolve away this material. Alternatively, or in addition thereto,physical methods, e.g., an autoclave, high pressure flushing with water,or others may be utilized to remove the crust of Cr-containing powder.

Next, an overlay coating may be applied to gas path surfaces of theblade 84 such as the airfoil 92 and the upper surfaces of the platform90 (step 216). The portions of the exterior may be along essentially theentire exterior or a portion of the exterior surface and gaspath-facingsurface(s) of the platform, the shroud, etc. The overlay coating asdefined herein includes, but is not limited to, cathodic Arc metallicbondcoat, and external duplex electron beam-physical vapor deposition(EB-PVD) ceramic coatings.

Detailed investigation of fielded components has shown that hotcorrosion damage commonly occurs on the internal surfaces in coolerlocations, such as towards the base of the airfoil and below theplatform, while the airfoil internals are commonly challenged in hightemperature oxidation further outboard towards the tip. By using thislocation-specific coating process, both degradation modes can beaddressed individually. Hot corrosion testing of the coating depositedin the relatively cooler locations of the airfoil internals shows nocoating penetration after 1000 hours of hot corrosion exposure at 1350 Fin a corrosive sulfate mixture (FIG. 7).

The use of the terms “a”, “an”, “the”, and similar references in thecontext of description (especially in the context of the followingclaims) are to be construed to cover both the singular and the plural,unless otherwise indicated herein or specifically contradicted bycontext. The modifier “ ” used in connection with a quantity isinclusive of the stated value and has the meaning dictated by thecontext (e.g., it includes the degree of error associated withmeasurement of the particular quantity). All ranges disclosed herein areinclusive of the endpoints, and the endpoints are independentlycombinable with each other. It should be appreciated that relativepositional terms such as “forward”, “aft”, “upper”, “lower”, “above”,“below” and the like are with reference to normal operational attitudeand should not be considered otherwise limiting.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reason,the appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A method of coating a component, comprising:aluminizing an array of internal passageways within the component; andchromizing a portion of the array of internal passageways within thecomponent after aluminizing the array of internal passageways.
 2. Themethod as recited in claim 1, wherein chromizing the portion of thearray of internal passageways comprises chromizing up to a demarcation.3. The method as recited in claim 2, wherein the demarcation is locatedat 20%-30% of an airfoil span of the component.
 4. The method as recitedin claim 2, wherein the demarcation is at 25% of an airfoil span of thecomponent.
 5. The method as recited in claim 2, wherein the demarcationis at an airfoil span location where local temperatures exceedapproximately 1600 degree F.
 6. The method as recited in claim 1,wherein chromizing the portion of the array of internal passagewayscomprises applying a chromium slurry into the array of internalpassageways within the component.
 7. The method as recited in claim 6,wherein the chromium slurry has a viscosity of 100-200 cp.
 8. The methodas recited in claim 1, further comprising applying a thermal barriercoat atop a bond coat on an external surface of the component.
 9. Acomponent, comprising: a substrate having an array of aluminizedinternal passageways within the component; and a chromium-enrichedcoating within a portion of the array of internal passageways.
 10. Thecomponent as recited in claim 9, wherein the substrate comprises asuperalloy.
 11. The component as recited in claim 9, wherein thechromium-enriched coating within the portion of the array of internalpassageways is located at an inboard section of the component.
 12. Thecomponent as recited in claim 9, wherein the chromium-enriched coatingwithin the portion of the array of internal passageways is located up toa demarcation that is located at 20%-30% of an airfoil span of thecomponent.
 13. The component as recited in claim 9, wherein thechromium-enriched coating within the portion of the array of internalpassageways is located up to a demarcation that is located at 25% of anairfoil span of the component.
 14. The component as recited in claim 9,wherein the chromium-enriched coating within the portion of the array ofinternal passageways is located up to an airfoil span location wherelocal temperatures exceed approximately 1600 degree F.
 15. The componentas recited in claim 9, wherein the chromium-enriched coating is 10-50microns thick.
 16. A component, comprising: an airfoil having an arrayof aluminized internal passageways, the array of aluminized internalpassageways chromized up to a demarcation.
 17. The component as recitedin claim 16, wherein the demarcation is located at 20%-30% of a span ofthe airfoil.
 18. The component as recited in claim 16, wherein thedemarcation is located at 25% of a span of the airfoil.
 19. Thecomponent as recited in claim 16, wherein the demarcation is located atan airfoil span where local temperatures exceed approximately 1600degree F.
 20. The component as recited in claim 16, wherein thecomponent is a turbine blade.